Passive control of gas turbine clearances using ceramic matrix composites inserts

ABSTRACT

A passive clearance control limits thermal expansion between stator components relative to rotor components. A control ring controls clearance in a passive manner and is located on or adjacent to stationary components which thermally expand during engine operation. The control ring is formed of material having low coefficient of thermal expansion such as CMCs (Ceramic Matrix Composites) and therefore limits, inhibits or restrains expansion of the adjacent stator components as temperatures increase. Limiting expansion of the stator component reduces rotor/stator clearances and limits parasitic leakage of fluid along the flowpath through the engine core.

BACKGROUND

Present embodiments relate generally to a gas turbine engine. Morespecifically, the present embodiments relate, but are not limited, toclearance control structures for stator components disposed opposite orat radially outward areas on rotating portions of the gas turbineengine.

A typical gas turbine engine generally possesses a forward end and anaft end with its several core or propulsion components positionedaxially therebetween. An air inlet or intake is located at a forward endof the engine. Moving toward the aft end, in order, the intake isfollowed by a fan, a compressor, a combustion chamber, and a turbine. Itwill be readily apparent from those skilled in the art that additionalcomponents may also be included in the engine, such as, for example,low-pressure and high-pressure compressors, and low-pressure andhigh-pressure turbines. This, however, is not an exhaustive list.

The compressor and turbine generally include rows of airfoils that arestacked axially in stages. Each stage includes a row ofcircumferentially spaced stator vanes and a row of rotor blades whichrotate about a center shaft or axis of the turbine engine. A multi-stagelow pressure turbine follows the multi-stage high pressure turbine andis typically joined by a second shaft to a fan disposed upstream fromthe compressor in a typical turbo fan aircraft engine configuration forpowering an aircraft in flight.

The stator is formed by a plurality of nozzle segments which are abuttedat circumferential ends to form a complete ring about the axis of thegas turbine engine. Each nozzle segment may comprise a single vane,commonly referred to as a singlet. Alternatively, a nozzle segment mayhave two vanes per segment, which are generally referred to as doublets.In a third embodiment, additional numbers of vanes may be disposed on asingle segment. In these embodiments, the vanes extend between an innerband and an outer band.

A typical gas turbine engine utilizes a high pressure turbine and lowpressure turbine to maximize extraction of energy from high temperaturecombustion gas. The turbine section typically has an internal shaftaxially disposed along a center longitudinal axis of the engine. Theblades are circumferentially distributed on a rotor causing rotation ofthe internal shaft. The internal shaft is connected to the rotor and theair compressor, such that the turbine provides a rotational input to theair compressor to drive the compressor blades. This powers thecompressor during operation and subsequently drives the turbine. As thecombustion gas flows downstream through the turbine stages, energy isextracted therefrom and the pressure of the combustion gas is reduced.

In operation, air is pressurized in a compressor and mixed with fuel ina combustor for generating hot combustion gases which flow downstreamthrough turbine stages. These turbine stages extract energy from thecombustion gases. A high pressure turbine first receives the hotcombustion gases from the combustor and includes a stator nozzleassembly directing the combustion gases downstream through a row of highpressure turbine rotor blades extending radially outwardly from asupporting rotor disk. The stator nozzles turn the hot combustion gas ina manner to maximize extraction at the adjacent downstream turbineblades. In a two stage turbine, a second stage stator nozzle assembly ispositioned downstream of the first stage blades followed in turn by arow of second stage rotor blades extending radially outwardly from asecond supporting rotor disk. The turbine converts the combustion gasenergy to mechanical energy.

During such operation of the gas turbine engine, it is desirable tominimize seal leakages or leakage between moving rotor components andthe opposed stator component radially outward of the rotor component.Limiting clearance in these areas improves performance of the engine.During operation, the large differences in the thermal and mechanicalgrowth of the rotating and stator components make it difficult to matchrotor/stator deflections. When relative deflections increase,rotor/stator clearances increase allowing leakage or increase parasiticflow. Transient mismatch also results in clearance opening duringtakeoff acceleration. Reducing engine air flow leakage results inimproved fuel efficiency and reduced fuel burn.

It would be desirable to overcome these and other deficiencies in orderto reduce clearances between rotor and stator components as well aslower exhaust gas temperature overshoot.

BRIEF DESCRIPTION OF THE INVENTION

According to some embodiments, a passive clearance control limitsthermal expansion between stator components relative to rotorcomponents. A low coefficient of thermal expansion passive clearancecontrol ring may be located on or adjacent to stationary componentswhich thermally expand during engine operation. The control ring isformed of material with low coefficient of thermal expansion such asCMCs (Ceramic Matrix Composites) and therefore inhibits or restrainsexpansion of the adjacent stator components as temperatures increase.Limiting expansion of the stator component reduces rotor/statorclearances and limits parasitic leakage of fluid along the flowpaththrough the engine core.

A clearance control ring assembly comprises a stator component disposedopposite a rotor component within a gas turbine engine, a clearancecontrol ring being formed of a single structure and extendingcircumferentially disposed radially outward of at least a portion of thestator component, the clearance control ring having a coefficient ofthermal expansion which is lower than the at least a portion of thestator component and, the at least a portion of the stator component islimited from thermal expansion and limited from growth from the rotorcomponent.

A clearance control ring assembly comprises a compressor dischargepressure seal stator, a honeycomb abradable opposite a compressordischarge pressure seal rotor, a ceramic matrix composite control ringextending circumferentially about the compressor discharge pressure sealstator, the ceramic matrix composite control ring being a one-piecestructure and, the ceramic matrix composite control ring limitingthermal expansion of the compressor discharge pressure seal stator andmaintaining tight seal clearance between the seal stator and the sealrotor.

All of the above outlined features are to be understood as exemplaryonly and many more features and objectives of the passive clearancecontrol in a gas turbine engine may be gleaned from the disclosureherein. Therefore, no limiting interpretation of this summary is to beunderstood without further reading of the entire specification, claims,and drawings included herewith.

BRIEF DESCRIPTION OF THE DRAWINGS

The above-mentioned and other features and advantages of these exemplaryembodiments, and the manner of attaining them, will become more apparentand the clearance control feature will be better understood by referenceto the following description of embodiments taken in conjunction withthe accompanying drawings, wherein:

FIG. 1 is a side section view of an exemplary gas turbine engine;

FIG. 2 is a side section view of a control ring near an exemplaryinterface at the compressor discharge pressure seal;

FIG. 3 is an exemplary control ring arrangement within a compressor;

FIG. 4 is a further exemplary embodiment of control rings;

FIG. 5 is a still further exemplary embodiment of a control ring;

FIG. 6 is a further embodiment with a retaining shoulder arrangement;

FIG. 7 is a further embodiment having a ring at radially outwardposition of case flanges;

FIG. 8 is a further embodiment having a plurality of control wraps;

FIG. 9 is a control ring embodiment having an angled surface forengagement with the casing;

FIG. 10 is a control ring with flanges to capture vane platforms and/orcase structures;

FIG. 11 is an embodiment of a movable control ring which closes a gap asan adjacent case structure expands;

FIG. 12 is an alternate embodiment of FIG. 11;

FIG. 13 is a control ring which serves as a blade shroud ring;

FIG. 14 is an alternate embodiment of a control ring serving as a bladeshroud ring;

FIG. 15 is a further alternate embodiment of a control ring and integralblade shroud ring;

FIG. 16 is side section view of an alternate embodiment having a biasingmechanism;

FIG. 17 is a partial isometric view of the embodiment of FIG. 16;

FIG. 18 is a side section view of a further alternate embodiment havingmultiple biasing mechanisms;

FIG. 19 is an isometric view of one embodiment of a biasing mechanism ofFIG. 18;

FIG. 20 is an isometric view of an alternate biasing mechanism of FIG.19; and,

FIG. 21 is an isometric view of the embodiment of FIG. 18 with schematicrepresentations of the biasing mechanisms of FIGS. 19 and 20.

DETAILED DESCRIPTION

Reference now will be made in detail to embodiments provided, one ormore examples of which are illustrated in the drawings. Each example isprovided by way of explanation, not limitation of the disclosedembodiments. In fact, it will be apparent to those skilled in the artthat various modifications and variations can be made in the presentembodiments without departing from the scope or spirit of thedisclosure. For instance, features illustrated or described as part ofone embodiment can be used with another embodiment to still yieldfurther embodiments. Thus it is intended that the present inventioncovers such modifications and variations as come within the scope of theappended claims and their equivalents.

Referring to FIGS. 1-21, various embodiments of a gas turbine engine aredepicted having a passive clearance control ring. The control ringinhibits thermal growth of itself and adjacent stator hardware topreclude increased spacing between a stator component and a rotorcomponent. As used herein the term stator component refers to astationary structure and the term rotor component refers to a rotatingcomponent rotating relative to the stator component.

As used herein, the terms “axial” or “axially” refer to a dimensionalong a longitudinal axis of an engine. The term “forward” used inconjunction with “axial” or “axially” refers to moving in a directiontoward the engine inlet, or a component being relatively closer to theengine inlet as compared to another component. The term “aft” used inconjunction with “axial” or “axially” refers to moving in a directiontoward the engine nozzle, or a component being relatively closer to theengine nozzle as compared to another component.

As used herein, the terms “radial” or “radially” refer to a dimensionextending between a center longitudinal axis of the engine and an outerengine circumference. The use of the terms “proximal” or “proximally,”either by themselves or in conjunction with the terms “radial” or“radially,” refers to moving in a direction toward the centerlongitudinal axis, or a component being relatively closer to the centerlongitudinal axis as compared to another component. The use of the terms“distal” or “distally,” either by themselves or in conjunction with theterms “radial” or “radially,” refers to moving in a direction toward theouter engine circumference, or a component being relatively closer tothe outer engine circumference as compared to another component. As usedherein, the terms “lateral” or “laterally” refer to a dimension that isperpendicular to both the axial and radial dimensions. The term “lowcoefficient of thermal expansion material” refers to a material whichgrows relatively less as the temperature increases.

All directional references (e.g., radial, axial, proximal, distal,upper, lower, upward, downward, left, right, lateral, front, back, top,bottom, above, below, vertical, horizontal, clockwise, counterclockwise)are only used for identification purposes to aid the reader'sunderstanding of the present invention, and do not create limitations,particularly as to the position, orientation, or use of the invention.Connection references (e.g., attached, coupled, connected, and joined)are to be construed broadly and may include intermediate members betweena collection of elements and relative movement between elements unlessotherwise indicated. As such, connection references do not necessarilyinfer that two elements are directly connected and in fixed relation toeach other. The exemplary drawings are for purposes of illustration onlyand the dimensions, positions, order and relative sizes reflected in thedrawings attached hereto may vary.

Referring initially to FIG. 1, a schematic side section view of a gasturbine engine 10 is shown. The function of the gas turbine engine is toextract energy from high pressure and temperature combustion gases andconvert the energy into mechanical energy for work. The gas turbineengine 10 has an engine inlet end 12 wherein air enters the core orpropulsor 13 which is defined generally by a compressor 14, a combustor16 and a multi-stage high pressure turbine 20. Collectively, thepropulsor 13 provides thrust or power during operation. The gas turbineengine 10 may be used for aviation, power generation, industrial, marineor the like.

In operation, air enters through the air inlet end 12 of the engine 10and moves through at least one stage of compression where the airpressure is increased and directed to the combustor 16. The compressedair is mixed with fuel and burned providing the hot combustion gas whichexits the combustor 16 toward the high pressure turbine 20. At the highpressure turbine 20, energy is extracted from the hot combustion gascausing rotation of turbine blades which in turn cause rotation of theshaft 24 about engine axis 26. The shaft 24 passes toward the front ofthe engine to continue rotation of the one or more compressor stages 14,a turbofan 18 or inlet fan blades, depending on the turbine design. Theturbofan 18 is connected by the shaft 28 to a low pressure turbine 21and creates thrust for the turbine engine 10. A low pressure turbine 21may also be utilized to extract further energy and power additionalcompressor stages.

Referring now to FIG. 2, a side section view of the interface betweenthe compressor 14 and the combustor 16 is depicted. Downstream of thecombustor 16, a first stage of the high pressure turbine 20 is shown aswell. In the depicted embodiment, a shaft 24 extends from the turbine 20to the compressor 14 such that rotation of the turbine 20 causesrotation of the high pressure rotor blades within the compressor 14.Along the shaft 24 is a compressor discharge pressure seal 30. The seal30 includes a stator component 32 and a rotor component 34. The rotorcomponent 34 of the instant embodiment is depicted as a labyrinth seal35 having a plurality of seal teeth 36 extending generally radiallyoutward. The stator component 32 is disposed opposite the rotorcomponent 34 and according to the instant embodiment, comprises at leastan abradable surface, such as a honeycomb material 38 and a support arm39. The honeycomb structure 38 is disposed upon a support arm 39. Sincethis picture is shown in section, one skilled in the art will understandthat the support arm 39 extends circumferentially about an engine centerline 26 (FIG. 1). Likewise, the abradable material 38 also extendscircumferentially. The circumferentially extending abradable material 38may be formed in multiple circumferential segments or a unitarystructure.

Disposed radially outward of the seal support 39 is a ceramic matrixcomposite control ring 40. The ceramic matrix composite control ring 40provides a low alpha material circumferentially extending about thestator component of the seal 30. According to some embodiments, thecontrol ring may be formed of ceramic matrix composite (CMCs). However,other materials may be utilized such as IN 909. The control ring 40typically has a lower coefficient of thermal expansion than othermaterials about which the control ring 40 is disposed. Therefore, thecontrol ring 40 limits thermal movement of the stator component orcomponents 32. The clearance control ring 40 extends 360 degrees alongan outer surface of the seal support arm 39. During normal operation,the seal support arm 39 and/or honeycomb material 38 expand radiallyoutward. The control ring 40 limits thermal growth of the statorcomponent 32 in the radially outward direction. This arrangementprovides for a tighter clearance between the rotor component 34 and thestator component 32 at operating temperature and condition where thermalexpansion normally increases clearance between rotor components 34 andstator components 32.

Referring now to FIG. 3, a side section view of a schematic highpressure compressor 14 is depicted in the instant embodiment. Thecontrol ring embodiment may be used alternatively to limit clearancebetween a rotating compressor blade 135 and an outer wall of thecompressor to limit air leakage. An exemplary compressor 14 includes anaft case 146 associated with a pair of rotor and stator components 134,132 which may be at any of various stages of the compressor 14. Theexemplary aft case 146 is formed of a first portion 146 and a secondportion 145. The first aft case 146 includes a blade shroud ring 149which defines at least a portion of the stator component 132 and isopposite the compressor blade 135. The blade shroud ring 149 defines anouter wall of the airflow path within the compressor 14. Beneath the aftcase 146 is a vane 133 and platforms 131. The vane 133 turns air in adesired manner to increase energy extraction at the adjacent downstreamblade 135. The upper platform 131 is received in part by portions of theaft case 146. For example, the platform 131 may include a flange, rib orfinger at axial ends to be received by grooves of the adjacent aft casemembers. Alternatively, the platform may have a groove which receives aflange or the like from the case 146. Other retaining embodiments may beutilized and these descriptions are not limiting.

The stator component 132 is depicted opposite a rotor component 134. Asdescribed in this embodiment, the stator component 132 is a blade ring149 which is formed by the aft case 146. Opposite the stator component132 is the rotor component 134, which according to the instantembodiment, is defined by the compressor blade 135. In order to inhibitthermal growth in a radially outward direction and limit clearanceincrease between the blade 135 and the blade shroud ring 149 a controlring 140 is disposed about the blade ring portion of the aft case 146.Positioned radially outward of the stator component 132 is a ceramicmatrix component control ring 140. The control ring 140 may be of aunitary or one piece circumferential structure and may have variouscross-sectional shapes. For example, according to the embodimentdepicted, the cross-section of the structure is generally rectangularand may have curved or sharp corners. However, this is merely exemplaryas will be seen, various shapes may be utilized. As with previousembodiments, the ceramic matrix composite control ring 140 limits thegrowth of the stator component 132 relative to the rotor component 134at operating temperature which limits leakage of air flow around therotor. In normal operation, the blade shroud ring 149 may generally growoutwardly away from the blade 135 due to thermal conditions. The instantcontrol ring 140 inhibits such growth beyond desired amount thuslimiting increased clearance. The control ring 140 is surrounded by anaft case 145 providing an upper support and the adjacent aft case 146providing a lower support for the control ring 140. Above the controlring 140, the aft cases 145, 146 may be bolted together.

Referring now to FIG. 4, a side section schematic view of an alternateembodiment of the high pressure compressor 14 is depicted. In thisembodiment, a rotor component 234 is again disposed opposite a statorcomponent 232. The stator component 232 is defined by a mid-case member247 which has a generally I-shaped cross-section and is connected to anddisposed between the aft case members 245, 246. The mid-case member 247extends circumferentially about the engine center line. The lowerportion of the mid-case member 247 is spaced from the aft case members245, 246 and defines a blade shroud ring 249. Beneath the blade shroudring 249 is a blade 235 which rotates about the engine center line. Aswith the previous embodiment, a platform 131 is disposed beneath aftcases 245, 246. Within this space between the blade shroud ring 249 andthe cases 245, 246 is at least one ceramic matrix composite controlrings 240, 241. According to the instant embodiment, two rings 240, 241are utilized—one on each side of the web of the mid-case member 247.These control rings 240, 241 extend circumferentially about the axis ofthe engine 10 and are formed of generally L-shaped structures to fitwithin the confined space between the mid-case 247, the aft cases 245,246 and vane platforms engaging the mid-case member 247 and the aftcases 245, 246. These rings 240, 241 inhibit thermal growth in radialoutward direction of the stator component 232 relative to the rotorcomponent 234. Thus, the function of limiting separation between thestator and rotor components provides improved sealing of air flow inthis region of the engine 10.

Referring now to FIG. 5, a further alternative embodiment is depictedwherein aft cases 345, 346 provide a space wherein a ceramic matrixcomposite control ring 340 is located. In this embodiment, the aft cases345, 346 are similar to those shown in FIG. 3. However, in thisembodiment, the control ring 340 is generally rectangular shaped withone corner partially removed. This allows for positioning of a vaneplatform 131. In this embodiment also, a lower side 341 of the controlring 340 is not parallel with the opposite long edge but instead extendsat an angle to the central axis of the engine 10. The control ring 340is again captured between the case 346 on a lower surface and the case345 on an upper surface and a portion of the adjacent vane platform 131.

Referring now to FIG. 6, an alternate embodiment of assembly FIG. 5 isdepicted in side section view. The aft case 445 includes a retainingshoulder 449. The control ring 440 includes a projection 443 whichengages the retaining shoulder 449 of the aft case 445. These structuresmay be reversed.

A second difference in the embodiment of FIG. 6 than that of FIG. 5 is agap 447 between the bottom surface of the control ring 440 and the uppersurface of aft case 446, which is positioned below the control ring 440.The gap 447 is provided between the ring 440 and the base ring of theaft case 446. The gap is used to delay contact between the shroud andcontrol ring until a specific engine cycle condition is reached. Bysetting this gap, and the point where contact occurs, deflections aretuned to meet minimum clearances across all operating conditions.Minimum attainable average clearances will result in maximum performancefor a given mission.

Referring to FIG. 7, an embodiment of the ceramic matrix control ring540 is depicted disposed about an upper interface or flange 541 betweenaft cases 546 and 545. The gap 547 is provided between the control ringand the aft cases so that contact is delayed. When the aft cases 545,546 are bolted together, the control ring 540 is captured therebetweenand limits radially outward growth of the cases 545, 546, such thatblade shroud ring 549 cannot expand, or expand beyond a desired amount,due to thermal expansion.

Referring now to FIG. 8, an alternate embodiment is depicted wherein aplurality of ceramic matrix composite wraps 640 are circumferentiallypositioned about the radially outer edge of at least one flange of oneor both of the aft cases 645, 646. The circumferential position of thewraps 640 inhibits thermal growth of the cases 645, 646. In turn thislimits clearance between the rotor component 634 and the statorcomponent 632 as previously described. The wraps 640 may include one ormore wraps which may be formed of various cross-sectional shapes. Thecross-sectional shapes may all be the same or may differ. In the instantembodiment, the wraps are circular in cross-section and are of the sameshape.

Referring now to FIG. 9, the aft cases 746, 745 include flangesextending radially relative to the engine axis and include a gap 747therebetween wherein a rib structure 748 of a control ring 740 extendsinto. One of the flanges of the cases 745, 746 may be at a slight anglerelative to the radial direction so that the control ring 740 isslightly wedged within the gap between the flanges of the cases 746,745. The control ring 740 also defines a blade shroud ring 749 definingstructure to maintain blade tip clearance.

Referring now to FIG. 10, an alternate embodiment is shown with aftcases 845, 846. A control ring 840 is interlocked between the cases 846,845 and receives flanges of adjacent veined platform hangers 851, 853.Again, with the instant embodiment, the control ring 840 defines theblade shroud ring 849 and therefore the blade tip clearance is limitedfrom expanding during operating conditions.

Referring now to FIGS. 11 and 12, views are depicted of aft cases 945,946 and 1045, 1046 having oppositely oriented control rings 940, 1040.The cases 946 and 1045 have angled surfaces 947, 1047 which engagecorrespondingly angled surfaces of control rings 940, 1040. As the casesgrow with temperature rise in an axial direction, the control rings 940,1040 slide downwardly to maintain or reduce the blade shroud ring toblade tip gap. Again, the desired function being to reduce parasitic airflow leakage. These embodiments also depict the control rings 940, 1040defining the blade shroud ring opposite the rotor component.

Referring now to FIG. 13, aft cases 1146, 1145 are positioned wherein acontrol ring 1140 is captured between axially extending flanges of eachof the cases 1145, 1146. The control ring 1140 defines the blade shroudring 1149 inhibiting thermal growth and leakage of air due to thermalexpansion.

Referring now to FIG. 14, an embodiment is depicted having aft cases1245, 1246 wherein the aft case 1246 is formed of ceramic matrixcomposite. Accordingly, the aft case 1246 defines the blade shroud ring1249 and the material inhibits growth as well as leakage between therotor and stator components 1234, 1232.

Referring now to FIG. 15, an embodiment is depicted wherein the aft case1346 has a flange 1347 formed of ceramic matrix composite. The adjacentblade shroud ring 1349 is formed of material with higher coefficient ofthermal expansion and grows more than the flange 1347. In operation, theflange 1347 functions as a growth limiter for the blade shroud ring1349. This assembly also limits increase in the blade tip gap at theblade shroud ring 1349 and inhibits leakage.

Since the low alpha materials grow less than the adjacent hardware, andthe need for an initial gap to maximize effectiveness, a biasingmechanism may optionally be utilized to keep the control ring positivelylocated (concentric with adjacent stator components) at all times andconditions. The biasing force may be provided by various forms ofsprings in order to provide bias in the radial direction, axialdirection or both. The biasing force maintains either or both of radialand axial constraint at shutdown, then deflects as the hardwarerequires. The spring also enables lower stresses in the CMC rings due todelayed engagement, which otherwise would prevent usage of low strengthCMC rings for clearance control.

Referring now to FIG. 16 a CMC ring 1440 is depicted between aft cases1445, 1446. A gap 1442 is located between the cases 1446, 1445. This gap1442 allows for growth of the CMC ring 1440 during operation of theengine. Located within the gap 1442 adjacent the CMC ring 1440 is abiasing mechanism 1444. The biasing mechanism 1444 is a spring accordingto the instant embodiment and provides an axial force and a radial forcethrough the L-shaped shape having first leg and second leg. Each legbiases in one of an axial direction and a radial direction. Each leg mayhave one or more projections 1447 to engage an adjacent structure andprovide a force on the ring 1440. The spring 1444 has a normal positionwhen the engine is in a cool condition to locate the ring 1440. Thespring 1444 inhibits the CMC ring 1440 from freely floating within thegap 1442.

Referring now to FIG. 17, an isometric view of a portion of thestructure 1444 is depicted. The biasing mechanism 1444 has first andsecond legs which are joined at an angle, for example about 90 degrees,however other shapes or angles may be utilized. Each of the legsincludes a one or more projections 1447 which engage adjacent surfacesand may work to provide spring force in whole or in part with the legsof the mechanism 1444. The biasing mechanism 1444 is shown engaging thering 1440 (FIG. 16) and may be a fully circumferential part or may beformed of segments which design a full circumferential structure orsegments which define a partial circular shape.

As opposed to the embodiment of FIG. 16 wherein the CMC ring is disposedcloser to the flowpath, a biasing mechanism may also be utilized in anembodiment wherein the CMC ring is disposed radially away from theflowpath. For example, referring again to FIG. 7, the CMC ring 540 islocated between flanges radially spaced from the flowpath. A biasingmechanism 1444 may be seated between the flanges and the CMC ring 540 topositively position the ring 540 relative to the flanges of cases 546,545.

Referring now to FIG. 18, a further embodiment is depicted withalternative biasing mechanisms. Disposed between cases 1546, 1545 is aCMC ring 1540, which includes at least one biasing mechanism 1544. Inthe instant embodiment, the CMC ring 1540 is biased in both axial andradial directions. The biasing mechanism 1544 is shown in perspectiveview in FIG. 19 and provides force in a radial direction. The biasingmechanism 1544 is a clip-like structure which wraps around the CMC ring1540, in the instant embodiment, by spreading the ends of the spring1544. Referring briefly to the perspective view of FIG. 21, the expanderspring 1544 is disposed within channels along the radial outer surfaceof the CMC ring 1540. In the instant embodiment, two springs 1544 areutilized to provide biasing force in the radial direction of the engine.

Referring now to FIGS. 18, 20 and 21, an additional biasing mechanism1543 is depicted which provides a biasing force in the axial directionof the engine. As with biasing mechanism 1544, the biasing mechanism1543 may take various forms. However, the instant embodiment utilizes aspring 1543 which is generally cylindrical in shape. The spring 1543further has a plurality of curved leaves with peaks and valleys whichengage one another and provide biasing force or compensating for thermalexpansion. In this way, the biasing mechanism 1543 positively positionsthe CMC ring 1540 but allows for some expansion in the axial direction.Similarly, the mechanism 1544 positively positions the CMC ring 1540 inthe radial direction but compensates for expansion of the ring 1540 andbiasing the ring 1540 once the engine cools. One skilled in the art willrealize that these embodiments may be used with any of the previouslydescribed embodiments in the instant specification and others that maynot be shown. Further, the biasing mechanisms 1543, 1544 may alsofunction in radial, axial or other directions if so oriented.

While multiple inventive embodiments have been described and illustratedherein, those of ordinary skill in the art will readily envision avariety of other means and/or structures for performing the functionand/or obtaining the results and/or one or more of the advantagesdescribed herein, and each of such variations and/or modifications isdeemed to be within the scope of the invent of embodiments describedherein. More generally, those skilled in the art will readily appreciatethat all parameters, dimensions, materials, and configurations describedherein are meant to be exemplary and that the actual parameters,dimensions, materials, and/or configurations will depend upon thespecific application or applications for which the inventive teachingsis/are used. Those skilled in the art will recognize, or be able toascertain using no more than routine experimentation, many equivalentsto the specific inventive embodiments described herein. It is,therefore, to be understood that the foregoing embodiments are presentedby way of example only and that, within the scope of the appended claimsand equivalents thereto, inventive embodiments may be practicedotherwise than as specifically described and claimed. Inventiveembodiments of the present disclosure are directed to each individualfeature, system, article, material, kit, and/or method described herein.In addition, any combination of two or more such features, systems,articles, materials, kits, and/or methods, if such features, systems,articles, materials, kits, and/or methods are not mutually inconsistent,is included within the inventive scope of the present disclosure.

Examples are used to disclose the embodiments, including the best mode,and also to enable any person skilled in the art to practice theapparatus and/or method, including making and using any devices orsystems and performing any incorporated methods. These examples are notintended to be exhaustive or to limit the disclosure to the precisesteps and/or forms disclosed, and many modifications and variations arepossible in light of the above teaching. Features described herein maybe combined in any combination. Steps of a method described herein maybe performed in any sequence that is physically possible.

All definitions, as defined and used herein, should be understood tocontrol over dictionary definitions, definitions in documentsincorporated by reference, and/or ordinary meanings of the definedterms. The indefinite articles “a” and “an,” as used herein in thespecification and in the claims, unless clearly indicated to thecontrary, should be understood to mean “at least one.” The phrase“and/or,” as used herein in the specification and in the claims, shouldbe understood to mean “either or both” of the elements so conjoined,i.e., elements that are conjunctively present in some cases anddisjunctively present in other cases.

It should also be understood that, unless clearly indicated to thecontrary, in any methods claimed herein that include more than one stepor act, the order of the steps or acts of the method is not necessarilylimited to the order in which the steps or acts of the method arerecited.

In the claims, as well as in the specification above, all transitionalphrases such as “comprising,” “including,” “carrying,” “having,”“containing,” “involving,” “holding,” “composed of,” and the like are tobe understood to be open-ended, i.e., to mean including but not limitedto. Only the transitional phrases “consisting of” and “consistingessentially of” shall be closed or semi-closed transitional phrases,respectively, as set forth in the United States Patent Office Manual ofPatent Examining Procedures, Section 2111.03.

What is claimed is:
 1. A clearance control ring assembly, comprising: astator component disposed opposite a rotor component within a gasturbine engine; a clearance control ring being formed of a singlestructure and extending circumferentially disposed radially outward ofat least a portion of said stator component, said clearance control ringhaving a coefficient of thermal expansion lower than a coefficient ofthermal expansion of said at least a portion of said stator component;and at least one spring disposed between the stator component and theclearance control ring, the at least one spring configured to provide aradial bias and an axial bias to the clearance control ring; whereinsaid at least a portion of said stator component is limited from thermalexpansion and limited from growth relative to said rotor component. 2.The clearance control ring assembly of claim 1, wherein said clearancecontrol ring is disposed in a compressor of the gas turbine engine. 3.The clearance control ring assembly of claim 2, wherein said statorcomponent includes a blade ring.
 4. The clearance control ring assemblyof claim 3, wherein said clearance control ring is disposed outwardlyand circumferentially about said blade ring.
 5. The clearance controlring assembly of claim 1, further comprising a mid-case ring disposedbetween a first aft case and a second aft case.
 6. The clearance controlring assembly of claim 5, wherein the first aft case and the second aftcase each has a flange.
 7. The clearance control ring assembly of claim6, wherein said clearance control ring is disposed between said flangesof the first aft case and the second aft case, respectively.
 8. Theclearance control ring assembly of claim 6, further comprising a gapbetween the flange of the first aft case and the flange of the secondaft case.
 9. The clearance control ring assembly of claim 1, wherein atleast one surface of said clearance control ring is angled relative toan axis of said gas turbine engine.
 10. The clearance control ringassembly of claim 1, wherein the at least one spring is L-shaped.